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I am new to the Aircraft Modification world. I have given a task to attach some Avionics equipment (only 2Kgs) to a small airplane (Twin Otter) fuselage, using appropriate mounting brackets.

The process I am following is

* Using Type certificate Found the Inertia requirements. ( Do I need to go through each FAR 23 amendment or use the Initial approved FAR 23 regulations? Di I need to consider existing running loads on Frames & stringers)
* Use the Inertia loading to find the mounting bracket bending , attachment fastener Shear .
(Since out of plane loads caused by bending due to the offset from Fuselage, can I use rivets? I read rivets are very bad in Tension, what could be the alternate fasteners?
* The fuselage skin is very thin (0.032") what could be the best option to stiffen the skin.
( Can I use doubler attaching to stinger )

Many Thanks


this is probably no the best way to approach this problem. many basic questions (normally would attach "equipment" with screws rather than rivets), not much definition ... where on fuselage ? how near to frames ? how big is the "equipment" ? (eg a small L-band antenna, a low profile GPS "puck", a large VHF blade ?? Inertia loads from FARs may not be relevant (but could work as "conservative" assumption). Who is certifying this installation (ie making the findings of compliance) ?

another day in paradise, or is paradise one day closer ?



Thanks for the reply,
The equipment is a sensor weigh 2kgs and mounting on fwd skin near to door ( centre of frames and stringer) and another at aft skin near empennage (close to stringer) . Our company is creating all the Technical data package and endorsed by a Part 21J approved agency. I am not fully understand the entire process, and confused by reading EASA and FAA documents .

Do you have any good example guidelines or reports ?.

Many Thanks



Inertia loads from FAR23 may not be conservative as GA is often gust critical. For something of such low mass, demonstration by structural test would be my first port of call for compliance to static loads (test equipment would consist of calibrated fish scales & a couple of zip ties). Edit: Brain fart, somehow I ended up thinking of an install with no fuse penetration.

Rivets in tension are not desirable but 2 kg shouldn't produce much in the way of rivet tension load (a single MS20479AD4 is good for ~80 lb+ tension).


attaching a 2kg mass to the plane should be easy ... for any inertial load. if it's low profile and, well, it is on a DHC6 then aeroloads should be small too. it's near a doorway ... maybe someone will stand on it ? again it is more common to attach things with screws (so they can be removed). if you use rivets the tension loads shouldn't be high.

If your company's is doing this STC then someone is writing the cert plan. who will make the findings of compliance ? It sounds as though your company doesn't have much experience/capability so maybe these are retained by the local certificating authority ? Whoever, this is the person to discuss how much detail they need to see.

another day in paradise, or is paradise one day closer ?


Is Your airframe mount point designed for high stiffness and minimal deflection... or was it arbitrarily selected for 'a convenient/other-component-factor'?

I imagine that a sensor install should probably be designed for high stiffness [all axis + torque] to minimize any strain-related movement of the component; and that external vibration/shock modes will be eliminated using suitable isolation/damping mounts. IF a design is stiffness driven, then loads should be almost inconsequential... unless it inadvertently becomes a functioning part of the airframe.

Is there a vibration design document for Your aircraft? That document may be a design driver. My acft has a document specifying vibration loads/modes that all tertiary/added equipment must be designed to withstand. These loads often exceed crash-loads for small components.

Also, as mentioned earlier... never forget potential for 'abuse' loads. A component mounted on a bracket in a 'hot-spot' may be a convenient hand-hold, armrest, target for nuisance-bumping, head-rest, etc.

A general rule of thumb is that flight critical structure should have secondary brackets installed by permanent [fatigue resistant] fastener installs whenever mounting temporary or high maintenance [high frequency remove-reinstall] parts. This isolates primary structure from any gross maintenance damage.

CAUTION. Electrical installs may also require adequate electrical bonding/grounding. Rivets in bare holes are great for general grounding... but high quality electrical bonding may require jumper-wires. Choose bonding/grounding methods carefully.

Regards, Wil Taylor

o Trust - But Verify!
o We believe to be true what we prefer to be true. [Unknown]
o For those who believe, no proof is required; for those who cannot believe, no proof is possible. [variation,Stuart Chase]
o Unfortunately, in science what You 'believe' is irrelevant. ["Orion", Homebuiltairplanes.com forum]


Excited to read everyone replies,

Yes I am including the gust inertia for analysis. Need to speak the supporting company for finding of Compliance

Presently Convenient Locations have choosed for Installation ( to Avoid existing avionics and structures)
Decided to put playcard to avoid Abuse loads

Decided to Mount bracket to Aircraft attachment using Rivets and a back plate (twice the thickness of skin, attached between the stinger ,Making use of existing stringer holes)

No Vibration design document available.. so not sure

Question remains:

* If I am attaching the back plate to Existing stringers or Frames, Do i need to consider any running loads.. If so which document should i refer?
* If the installation location fall between a stringer , Can we cut the stringers to make room?

* Also there is no CAD drawing available for Twin Otter Fuselage surface, How do I make the exact skin profile in CAD??
* DO I need to Assess any Drag ??


You have to account for the lost area from the skin. Normally for a simple install I would suggest starting with a SRM repair scheme for a hole in the skin at the applicable location and modify to accommodate the antenna (taking care not to invalidate any of the underlying assumptions the OEM used, shouldn't be hard with a DHC-6 but somewhat harder on more complex aircraft) or start from first principles and calculate load capacity in the skin joint at the nearest edge of the fuse skin (count the rivets per inch& ID the rivet type etc) to get the lost tensile capacity.

A placard wont stop all abuse loads, it will just mean the installation ends up being considered unreliable.


"Decided to put playcard to avoid Abuse loads" ... good project decision, bad real world decision. One advantage you get from abuse loads is creating a structure to take large loads ('cause the "real" loads (aero or inertia) are some small) and to create a structure fix enough to avoid vibration issues. Be very careful if you're anywhere near the pilot step.

"Decided to Mount bracket to Aircraft attachment using Rivets and a back plate (twice the thickness of skin, attached between the stinger ,Making use of existing stringer holes)" ... I think this is a reasonable/good plan ... a permanently attached adapter plate which the equipment screwed into for easy removal. It also allows the adapter plate to be curved on one side, to interface nicely with the fuselage, and flat on the other, for the equipment.

If your equipment doesn't have a vibration spec then it has no requirements. However, a requirement could be buried, eg in how they tested for DO160 ... maybe they mounted the equipment to a rigid base, so attaching to a "non-rigid" fuselage may invalidate this ?

as noted above, you need to account for removed area. The SRM is a guide. As for preexisting airplane loads, try to avoid cutting stringers ... if you Have to, then replace with two stringers on either side of the installation (cut-out?), a full bay so reasonably longer than the cut-out. A much more involved installation. Loads ... you Could assume the full Ftu on the cut area and design the adapter riveting to suit ... be very careful with this grossly conservative assumption as it creates very high loads, which leads to extensive riveting.

Drag ... it is a DHC6 ... Vmo is what 170 kts ? If the equipment is low profile then a good story for neglecting drag loads. If not and you calc drag loads as small, be aware of the "flutter and vibration" requirements (23.251?). Mind you DHC6 is an old design, CAR3 and SFAR23 ... check the type certificate.

Talk early and often with whoever is making findings of compliance ... they are the real guide/driver to what you need to do. If they're satisfied then they'll sign; if not, then no sign ! Yes, you are held hostage by them !!

another day in paradise, or is paradise one day closer ?


If you plan on putting a single layer thick (read: stiff) backing plate on the thin skin, this is a poor fatigue detail. You have said this is Part 23 but there are still damage tolerance requirements (23.573, 23.574). At the very least your STC will be maintaining the Type Cert basis so your added structure needs to comply with those parts. Since this is an STC you will need to complete this evaluation at the time of certification.

Keep em' Flying
//Fight Corrosion!


basis of cert of DHC6 pre-dates DT (it barely made fatigue). Unpressurised fuselage skin is unlikely to be fatigue critical.
I think that even the current S400 model only picked up new rules as they directly applied to the instrumentation changes, and most of the aircraft is grand-fathered (or great-grand-fathered). The wing, lift strut, and tension carry-thru structure are the most fatigue critical items.

another day in paradise, or is paradise one day closer ?


Agreed. I was just trying to say that if this is an STC, the original cert basis is just the bare minimum. It sounds like the OP doesn't know yet what the STC certification basis will be. If for some reason they decide to show something different, there may be different analytical requirements. Just something to be aware of is all.

Cert. basis per TCDS A9EA is actually CAR 3 Amdt. 3-1 to 3-8. Suggest the OP read through Amdt 3-2. If the fatigue life is okay because the structure is non-pressurized, fine. Just saying it is important to be aware of the regulatory basis being shown, and not to ignore anything, even if it is trivial.

Keep em' Flying
//Fight Corrosion!


Wouldnt it be a good idea to incorporate skin mapping into your design as part of the location selection process?



another day in paradise, or is paradise one day closer ?


Thanks for all suggestions

Vmo is 170 kts. The drag force calculated for the profile is less.But how do we calculate the Total drag for comparison?.

We are communicating with the people who making findings of compliance. No immediate responses..

Following this assumption.. "Unpressurised fuselage skin is unlikely to be fatigue critical"

Also I have no clue about F&DT. Which is the good starting point ?. Do we need any software tools needed?. Is there any example Fatigue calculation available to get an Idea?

Many Thanks


you're not interested so much in total airplane drag, but rather the drap on the equipment (so to size the attachment of the equipment). hint ... drag will be small ... design for abuse loads.

F/DT ... is there a fatigue life on the part you're mounting on ? (probably not ... there are a few life limited parts on the fuselage (refer PSM 1-6-7A) ... stay away from them. Then the question becomes is your installation creating a fatigue load (a buffet ??) ...

another day in paradise, or is paradise one day closer ?

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